This invention relates to combustors in turbo machinery and specifically, to the cooling of combustor liners in gas turbine combustors.
Conventional gas turbine combustion systems employ multiple combustor assemblies to achieve reliable and efficient turbine operation. Each combustor assembly includes a cylindrical liner, a fuel injection system, and a transition piece that guides the flow of the hot gas from the combustor to the inlet of the turbine. Generally, a portion of the compressor discharge air is used to cool the combustion liner and is then introduced into the combustor reaction zone to be mixed with the fuel and burned.
In systems incorporating impingement cooled transition pieces, a hollow sleeve surrounds the transition piece, and the sleeve wall is perforated so that compressor discharge air will flow through the cooling apertures in the sleeve wall and impinge upon (and thus cool) the transition piece. This cooling air then flows along an annulus between the sleeve surrounding the transition piece, and the transition piece itself. This so-called xe2x80x9ccross flowxe2x80x9d eventually flows into another annulus between the combustion liner and a surrounding flow sleeve. The flow sleeve is also formed with several rows of cooling holes around its circumference, the first row located adjacent a mounting flange where the flow sleeve joins to the outer sleeve of the transition piece. The cross flow is perpendicular to impingement cooling air flowing through the holes in the flow sleeve toward the combustor liner surface.
The presence of this crossflow has a direct influence on the cooling effectiveness in the zone near where the first row of jets in the flow sleeve would have been expected to impingement cool the combustor liner. Specifically, the crossflow impacts the first row of flow sleeve jets, bending them over and degrading their ability to impinge upon the liner. In one prior design of the flow sleeve impingement jets there are three rows of 24 jets spaced evenly around the circumference of the flow sleeve. This jet pattern in the presence of the strong crossflow from the transition piece impingement sleeve produces very low heat transfer rates on the liner surface near the first row of jets. This low heat transfer rate can lead to high liner surface temperatures and ultimately loss of strength. Several potential failure modes due to the high temperature of the liner include, but are not limited to, cracking of the aft sleeve weld line, bulging and triangulation. These mechanisms shorten the life of the liner, requiring replacement of the part prematurely.
This invention enhances the cooling of the liner in Dry Low NOx type gas turbine combustors where jet impingement is used to cool the aft portion of the combustor liner. Even though there is a strong crossflow resulting from the transition piece cooling flow, the negative impact of the crossflow is minimized by the use of collars or cooling conduits, also referred to as a thimbles, that are inserted into the cooling holes in the combustor liner flow sleeve, through which the cooling jets pass. These thimbles provide a physical blockage to the cross flow which forces the crossflow into the desired flow path while simultaneously ensuring that the cooling jets effectively impinge on the combustor liner surface to be cooled.
The thimbles or collars are preferably mounted in each hole of at least the first row of holes at the aft end of the flow sleeve, adjacent a mounting flange where the combustor liner and transition piece are joined. This arrangement decreases the gap between the jet orifice and impingement surface; blocks the cross flow that deflects the jets and forces it into the desired flowpath for the subsequent jet rows; allows the diameter of the jet to be smaller and thereby reduce cooling air; and provides consistent and accurate control over the location of jet impingement. It also stabilizes unwanted axial oscillation of the first row of jets, and prevents the formation of a thick boundary layer (and resulting reduced heat transfer) upstream of the first row of jets.
Accordingly, in its broader aspects, this invention relates to a combustor for a turbine that includes: a combustor liner; a first flow sleeve surrounding the liner, the flow sleeve having a plurality of rows of cooling holes formed about a circumference of the flow sleeve; a transition piece connected to the combustor liner and the flow sleeve and adapted to carry hot combustion gases to a stage of the turbine; the transition piece surrounded by a second flow sleeve; wherein a first of the plurality of rows of cooling in the first flow sleeve is located adjacent the transition piece; and further wherein one or more cooling conduits are mounted in the cooling holes of at least the first of the plurality of rows of cooling holes.
In another aspect, the invention relates to combustion liner flow sleeve adapted for mounting in surrounding relationship to a combustion liner, the flow sleeve comprising a tubular body formed with plural rows of cooling holes, one of the plural rows located adjacent one end of the flow sleeve; and a cooling conduit mounted in each hole of at least the first row of holes, the cooling conduits projecting radially into the flow sleeve.
In still another aspect, the invention relates to a method of cooling a combustor liner of a gas turbine combustor, the combustor liner having a substantially circular cross-section, and a flow sleeve surrounding the liner in substantially concentric relationship therewith creating an annulus therebetween; the method comprising a) providing a plurality of axially spaced rows of cooling holes in the flow sleeve, each row extending circumferentially of the flow sleeve, one of the rows adjacent a forward or downstream end of the flow sleeve; b) locating a plurality of cooling conduits in the cooling holes of at least the first of the rows, the cooling conduits extending radially towards, but not engaging, the liner; and c) supplying cooling air to the annulus such that the cooling conduits direct the cooling air against the liner.